Fifth-Generation Target Drone Phase I Design

Transcription

1 11th AIAA Aviation Technology, Integration, and Operations (ATIO) Conference, including the AIA September 2011, Virginia Beach, VA AIAA Fifth-Generation Target Drone Phase I Design Dennis Carter 1 Air Force Research Laboratory Patrick Burris 2 Applied Resources, Inc. Steven Brandt 3 United States Air Force Academy The Fifth Generation Target Study (5GTS) was a conceptual design task for a large target drone aircraft capable of representing key 5th generation fighter signature and performance attributes. The target drone s intended uses are to test air-to-air and surfaceto-air missiles, and to ensure that modern tracking systems are capable of identifying and targeting 5th generation fighters. The costs associated with a new military design were unaffordable for a target, so the team looked into taking advantage of the materials and methods that were developed for the new Very Light Jets in general aviation. This paper is the second in a series of four on the conceptual design of 5GTS. It describes six conceptual designs created by Sierra Technical Services, the Air Force Research Laboratory, the Air Force Aeronautical Systems Center, and the United States Air Force Academy. The six designs were analyzed and evaluated relative to the evolving 5GTS requirements. These analyses helped the design team understand and refine the 5GTS requirements and also eliminated three of the six candidate configurations from further consideration. Nomenclature 5GTS = 5 th Generation Target AoA = Analysis of Alternatives A005, A008, D300, L005 = Configuration Names ASC = Aeronautical Systems Center CS = Clean Sheet Design FSAT = Full Scale Aerial Target (early program name) FMA = Foreign Military Aircraft M&S = Modeling and Simulation MRAAM = Medium Range Air-to-Air Missile OSD = Office of the Secretary of Defense P k = Probability of Kill QF-16 = Droned Version of F-16 RCS = Radar Cross Section RMA = Retired Military Aircraft SAR = Specific Air Range, miles flown per pounds of fuel burned SRAAM = Short Range Air-to-Air Missile STS = Sierra Technical Services VLJ = Very Light Jets T I. Introduction HE 5th Generation Target Study (5GTS) was a conceptual design task for a large target drone aircraft capable of representing key 5th generation fighter signature and performance attributes. The target drone s intended uses are to test air-to-air and surface-to-air missiles, and to ensure that modern tracking systems are capable of identifying and targeting 5th generation fighters. Aerial target systems are aircraft that are utilized in live fire testing to validate the effectiveness of a weapon system. Their purpose is to replicate the characteristics of threat aircraft, such as performance, signatures, and countermeasures. This testing is stipulated by Public Law, Title 10, Section 2366 of the U.S. Code, which requires that all new/improved weapon systems demonstrate their lethality prior to 1 Engineer, Air Vehicles Directorate, AFRL/RB, Wright-Patterson AFB, OH 45433, AIAA Associate Fellow 2 Project Engineer, Applied Resources, Inc., 1525 Perimeter Parkway, Suite 210, Huntsville, AL Professor of Aeronautics, Department of Aeronautics, Suite 6H156, USAF Academy, CO, 80840, AIAA Member This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

2 being approved for production. In addition, testing must be conducted to ascertain effective launch envelope and to demonstrate safe weapon separation from the parent aircraft. The U.S. military has used aerial targets for over thirty years to improve the lethality of their weapons systems. A host of retired fighter aircraft has served at targets for this program. The aircraft now used is the venerable F-4, but the number of airframes suitable for economical conversion is dwindling quickly and they are projected to be all expended by An interim solution is to convert the F-16, until a target more representative of the 5th Generation threat can be developed. Those attributes that characterize a 5th Generation fighter are shown in Table 1. Table 1. Characteristics of a 5 th Generation Fighter Significantly lower signatures in both radar cross section and infrared Low probability of intercept sensors Integrated electronic warfare Supermaneuverability Advanced materials Internal weapons Supercruise II. Background The Analysis of Alternatives (AoA) study, discussed in the previous paper 1, after investigating many options arrived at the conclusion that the only viable airframe on which to base an interim target drone was the F-16. It was understood at the time that it would be at best an interim answer to fill the gap between the depletion of the existing QF-4 inventory and the development of a truly representative 5 th Generation target. The design work for the 5 th Gen began in the summer of 2007 by the Office of the Secretary of Defense (OSD), under the Director of Operational Test and Evaluation. The initial effort with the project was to ascertain what the using community was looking for in a new target drone. This was accomplished in a requirements conference held for the advanced munitions program offices, the target operating organizations, and the tactics development offices. The results from this conference demonstrated that two options for the new target existed, one that was a replacement to the QF-16 and the other as a complement to the QF-16. These two sets of requirements are shown in Table 2. As the Complementary design showed the potential of being significantly less expensive than a complete replacement, both were carried forward into the design studies, even though the Complementary design would require the continuation of the QF-16 program. While the performance of the Complementary design would be significantly reduced from that of the Replacement design, analysis of historic data (Figures 1 & 2) indicated that the stipulated performance would encompass roughly 90% of the testing of the previous twenty years. Even with the assumed migration of testing to higher performance levels, it was felt that the Complementary design could still fulfill approximately 80% of the testing for the next fifteen years, with the QF-16 completing the remainder of the required testing. In parallel with the user workshop, a survey was conducted of the aircraft industry to determine the best approach to manufacturing the drone. Discussions with the major aircraft manufacturers demonstrated that developing an expendable aircraft through them would be unaffordable. This led to discussions with the General Aviation industry, which was developing affordable aircraft for the Very Light Jet market, like the HondaJet VLJ shown in Figure 4a. These companies were most eager to participate, and they were confident that they could produce the 5th Gen target in the schedule that was desired and under our cost goal. They expressed concern that we were looking for performance outside their normal flight envelope, but felt that this would not be a serious problem. A larger problem was that they had no experience in designing for reduced signatures as well as the attendant security requirements of extremely low signature levels. They recommended that the Government generate the outer mould lines of the configuration and they would develop the internal structural design and subsystem layout, based upon Government requirements. This recommendation led initially to two Air Force design studies. The first study was with Air Force Research Laboratory (AFRL) at Wright-Patterson AFB, OH. The second was with the Department of Aeronautics at the United States Air Force Academy, in Colorado, which used the project for a design study for the faculty, but also as

3 a series of academic exercises for the cadet design classes. Later a group of retired Lockheed engineers were placed on contract to give an industry look at the design problem. The common mission for all configurations is shown in Figure 3. Table 2. 5th Gen Target Requirements Design Goals Parameter Complimentary Replacement Operating Altitude Minimum, ft Maximum, ft Airspeed kft KCAS 36 kft Maneuverability Instantaneous 15 kft 9 10 Sustained 15 kft 3 7 Ps at 6 g's No reqm't Match F-16 Roll rate Match F-16 Match F-16 Physical Size Length, ft Wing span, ft Wing area, sq. ft Radar Signature RCS (nose-on), sqm Good Better RCS (beam), sqm Reduced Reduced External payloads Maximum number 2 3 Weight (per station), lbs 400 lbs 1000 lbs AI radar Weight, lbs 200 lbs 300 lbs Internal EA payload 1 pod 2 pods Endurance Minimum, hrs III. AFRL-led Design Studies Due to manpower limitations, the AFRL-led studies were split. AFRL Air Vehicles Directorate (AFRL/RB) designed a complementary target configuration and the Air Force Aeronautical Systems Center s Advanced Design Group (ASC/XR) was asked to investigate a replacement configuration. AFRL provided overall leadership of the split effort. A. Complementary Study The AFRL design process was initiated by calibrating analysis codes to two VLJ configurations that were currently undergoing flight testing. This calibration was necessary for these vehicles construction more closely resembled that anticipated in this study, rather than the conventional military aircraft to which the codes had been previously calibrated. The VLJs selected were the HondaJet (Figure 4a) and the Eclipse 500 (Figure 4b) by virtue of the amount of geometry and performance information that was available on the Internet.

4 A sizing trade study was begun to ascertain the gross vehicle parameters that would fulfill the mission shown in Figure 3. Utilizing typical fighter values for the design parameters as shown in Table 3, the capabilities of each combination were calculated using the NASA/Langley Flight Optimization System (FLOPS) analysis program. Baseline Mission Requirements 15 min taxi/warm-up time Distance credit allowed for climbs No distance credit for descents In area 20K 60min(Threshold)/90min(Objective) 30 min 10K (fuel reserve) Takeoff 100 nm 100 nm One of the following (as appropriate): -5 min Max Pwr@ 20K & 1g - 5 min & 1 g w/ one 6g circle - 5 min 20K & 1g w/ one 1.5M pass at 50K for 60 nm Land 5 Figure 3 Baseline Mission

5 Table 3. Design Space Variable Symbol Lower Bound Upper Bound Thrust-to-Weight Ratio T/W Wing Loading W/S 40 psf 50 psf Aspect Ratio AR FLOPS design point results for the constraints of Takeoff Field Length, Cruise Mach Number, Sustained 3G Turn, Acceleration from 0.5M to 0.95M, Landing Field Length, Maximum Mach Number, and Climb Rate are plotted on Figure 5, with the goal of minimizing empty weight while satisfying the constraints. The optimum configuration would have a W/S of 45 lbs/ft 2 and a T/W of Since all aircraft experience a weight growth through its development, a T/W of 0.8 was chosen. As the plot shows, a nominal increase in W/S would still be a viable design. Based upon these parameters, the specified 800-lb payload to be carried either internally or externally, the design mission and available engines, an initial design for the Complementary Configuration is shown in Figure 6. The AFRL twin-engine concept was shaped to meet the signature requirements shown in Table 2. A notional internal subsystem layout is presented in Figure 7. The serpentine inlets result in 100% line of sight blockage of the front faces of the engines. The fuel would be carried in the wing and in the fuselage (green). The equipment bay location (yellow) would permit access without bending or hard stands. The maneuverability of the AFRL twin-engine concept is shown in Figure 8. It would cover more than 95% of the historical shot matrix at a significantly lower cost than previous drones. The AFRL twin-engine design met the design requirements levied by OSD and the subject matter experts. It was presented to OSD at the mid-term review held at the Pentagon in May While the design satisfied the requirements, it was felt that the configuration would be too expensive to manufacture in production lots based upon the cost estimates developed independently for OSD. AFRL was directed to generate a smaller, single engine design that might be affordable.

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7 Figure 8. AFRL Twin Maneuverability

8 This study resulted in a minimal configuration to satisfy the modified requirements. The external carriage requirement was removed, but the internal payload remained at four hundred pounds, otherwise the mission and requirements remained the same. Figure 9 shows the result of that study. Discussions with Williams International revealed that they were developing a new version of their FJ44 engine, the Dash Five (-5). This engine, while more expensive than the -4s used in the twin engine design, promised a sea level static thrust of 5,500-lbs, significantly more than the 3,600-lbs from the -4. This increase in thrust, coupled with the reduction in vehicle size attendant with removal of one engine, promised performance comparable to the original twin-engine configuration. As shown in Figure 9, the new configuration was substantially smaller than the previous AFRL one. With a 26% reduction in wing area and reducing the empty weight by a third, the overall takeoff gross weight was cut almost in half for the same mission. This result was certainly in the desired direction of producing a substantially smaller and less expensive concept. The internal arrangement for the single engine design shown in Figure 10 was quite similar to the earlier twinengine configuration. The bifurcated inlet helped to hide the engine s front face reducing the front sector RCS signature. The engine nozzle was recessed well forward of the trailing edge of the tail. This was an attempt to hide the hot parts of the engine for passive infrared reduction, as well as reducing the angle at which the engine s rear face could be observed by radar. These measures, coupled with the obvious parallel edges and lack of corner reflectors, resulted in an acceptable level of radar signature reduction. Analysis of the new configuration, with its reduced wing loading and increased thrust to weight ratio, led to significantly improved performance when compared to AFRL Twin-Engine configuration. Maximum sustained turn rate at 0.8M was 20% better, and at 0.4M was over 100% improved.

9 . B. Replacement Configuration The target requirements as listed in Table 2 established a lofty goal for performance of a new-build target drone that could fully replace the F-16. Figure 11 presents the salient points as they impact the design. Because the entire AFRL/RB design team was concentrating on the Complementary configuration, a team from the Aeronautical Systems Center (ASC/XR) was asked to support the project through the development of a configuration that would have performance comparable with the F-16, with a significantly lower RCS signature. This support was most willingly given and the ASC team produced a credible design in very short order. While the visual signature requirement of being at least the size of an F-16 was significant, the primary driver of this configuration study was that of an available engine. While the afterburning F404 might have been a better choice, they were not available in a large number. And since we were constrained to only those engines that were in sufficient quantity to support the entire drone production run if this configuration were selected for series production, this limited our selection to the F100 pulled from retired F-16 aircraft. As the F-16 was originally designed as a minimal fighter, it soon became apparent that the new vehicle would weigh about the same as the F-16. The results of the sizing trade study are shown in Figure 12. The Best Case configuration for a QF-16 replacement was the result of an additional sizing iteration on a baseline concept that was deemed too heavy to acceptably fulfill the mission. This Best Case design is shown in Figure 13, along with its internal layout in Figure 14. The weight statement for the Best Case is shown in Table 4. Even though Best Case was intended to minimize weight, it still was much heavier than the other conceptual designs considered, and coincidently almost exactly the empty weight of the YF-16 prototype. The shear size and bulk of the F100 engine made further weight reduction impractical.

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12 Table 4. Replacement Weight Statement Item Best Case Weight (lbs) TOGW Structure 5,882 Propulsion 3,969 Fixed Equipment 2,898 Empty Weight (We) 12,749 Operational Items 277 Operational Weight Empty (OWE) 13,026 Fuel 6,974 Payload 400 IV. USAFA Design Study In an effort to reduce cost and improve level-flight performance, the USAFA investigated several delta-wing versions of a baseline configuration designated A005. The A005 concept was developed by Sierra Technical Services, a contractor hired to provide the 5GTS design team with the necessary staff of experienced engineers in a wider variety of specialties. A005 was characterized by a slender chined nose, twin scarfed and chined inlets, diamond wing and horizontal stabilators, and twin canted verticals. A more refined version of this concept, designated A008 will be described later in this paper. See Figure 23 for more details of this configuration. Each delta wing configuration retained the baseline A005 fuselage, inlets, canted twin vertical tails, and initially, twin FJ44-3 engines. All delta-wing configurations studied deleted the baseline A005 stabilators and associated rear-fuselage structure. Deletion of these structures combined with the well-known superior structural efficiency of delta wings produced significant weight and cost savings for all delta configurations evaluated. The 5GTS design team decided to limit the delta wing area to no less than 300 square feet. With this limit, the USAFA design team optimized the A005 Delta configuration for maximum level-flight performance while retaining maneuverability equivalent to that of the baseline A005. For this process, maneuverability equivalence was defined as enclosing the same number of historical 5K to 15K altitude turn points. With these criteria, the USAFA optimization software identified a 57-degree-sweep, AR=1.8 delta wing as the optimum wing planform. However, shifting to a slightly lower AR = 1.6 produced a wing with trailing edges aligned (for RCS reasons) with the A005 baseline and with performance nearly equal to the optimum. This 57- degree-sweep, AR=1.6 delta wing and a table of descriptive parameters for this design are shown in Figure 15. Note that the figure shows the delta wing configuration, labeled D300 to denote its 300 square foot delta wing, with engine inlets that are flush with the leading edge of the wing. These inlets could be extended forward so that their upper surfaces and sharply chined outer edges would form strakes that may enhance the aircraft s maneuverability. Water- and wind-tunnel testing was conducted to determine if this is desirable as part of Phase II, which is described in the third paper of this series. Performance estimates for delta-winged aircraft must take into account factors not usually included in estimates for conventional aircraft configurations. These factors include vortex lift and drag associated with leading-edge vortices as well as additional trim drag associated with the lack of a separate trimming surface. For a delta wing aircraft with positive static longitudinal stability, the magnitudes of these two types of drag can be very significant. In order to validate and calibrate performance estimation methods for this study, the USAFA team first modeled a well-known statically-stable delta-wing fighter aircraft, the F-106. By adjusting the analysis to use actual installed values for engine thrust and to increase drag-due-to-lift by 30%, the team achieved good agreement between analysis and actual performance for the F-106. Figure 16 shows the agreement between predicted and actual turn performance for the F-106 at 10,000 ft altitude. In the figure, the dotted green line represents the actual sustained turn or Ps = 0 capability of the aircraft. The red line labeled Ps = 0 is the same capability predicted by the corrected analysis. The good agreement between predicted and actual over a wide range of Mach numbers is obvious.

13 Figure 15. D300 - A005 Fuselage with 300 sq ft Delta Wing Figure 16. F-106 Predicted and Actual Maneuverability Diagram

14 The team then used the same corrections to predict the performance of the statically-stable delta-winged D300. They compared these with un-corrected predictions for the baseline A005 configuration. This was very generous to the A005, because it certainly experiences some vortex drag from chine and inlet/strake vortices. The A005 also undoubtedly has a non-parabolic drag polar at high values of lift coefficient. Ignoring these two factors made the A005 look better in the comparison charts then it likely would in reality. Figure 17 shows Ps = 0 contours for the A005, the delta-winged D300, and the F-16 superimposed on a plot of historical data for previous target drone usage. Note that the D300 encloses 34 more historical points in its sustained level-flight envelope than does the A005. Both, however, fall well-short of the F-16 s capability. Figure 17. Baseline and Delta Wing Drones and F-16C Sustained Level-Flight Envelopes Figure 18 shows sustained turn data for the same aircraft and history. On this diagram A005 and D300 both enclose approximately the same number of historical points in their sustained envelopes. Once again, both fall well short of the F-16 s capability. In an effort to remedy the shortfall in FJ44-powered drone sustained performance, the USAFA team studied an alternative engine. The Honeywell/International Turbine Engine Corporation (ITEC) F124 low-bypass-ratio turbofan powers the X-45 and its afterburning brother, the F125 powers the Republic of China s (Taiwan) Indigenous Design Fighter. When the USAFA team substituted a single F125 for the two FJ44s in the D300, performance improved dramatically. Figures 19 and 20 show that with the F125, the D300 becomes essentially equivalent to the F-16 in terms of sustained envelopes, and vastly superior to the FJ44-powered drones. Analysis suggested that the empty weight savings for the D300 delta over the A005 baseline could be on the order of 1,200 lbs. This in turn would produce a cost savings on the order of $1M per aircraft. The delta-winged aircraft would also be easier to disassemble, containerize, decant, and re-assemble, because only two wing panels would need to be removed and re-installed.

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16 Figure 20. F125-Powered Delta Wing Drone Sustained Turn Compared with F-16 The USAFA team postulated that if the cost savings were instead used to buy the more expensive F125 afterburning turbofan engine for the drone, the result would be an aircraft which would cost $1M to $2M more than the FJ44-powered A005, but which would be vastly more capable. This made the F125-delta combination even more appealing. The USAFA team created water- and wind tunnel models to determine stability characteristics of the delta-wing drone configuration. The results were as expected. The delta-winged models exhibited excellent stability characteristics out to 18 degrees angle of attack. Overall, the FJ44-powered 300-sq-ft delta wing configuration offered lower costs, superior level-flight performance, and equivalent but not identical sustained turn performance to the baseline A005. It also had superior stability characteristics. Coupled with the more-expensive F125 afterburning engine, the delta-winged drone would be equivalent to and a very capable replacement for the QF-16. Unfortunately, the extra cost of the F125 compared to the FJ44 made replacing the QF-16 unattractive. Instead, using QF-16s for the high-performance edges of the testing envelopes and using a lower-performing complementary stealthy drone for the hearts of the envelopes offered the best value-to-cost ratio. For this reason, the delta-wing study suggested eliminating the QF-16 replacement concepts from consideration and focusing on complementary concepts only. V. Sierra Technical Services Complementary Configuration Sierra Technical Services (STS) was tasked to design a complementary solution for the 5 th Generation Target Study. Their efforts culminated in the A008 configuration, an aircraft sized to meet the requirements for a complementary solution including the physical size requirement (same planform area as an F-16). As a cost trade investigation, STS was also asked to estimate how much the A008 configuration could be shrunk if the primary performance requirements were retained, but the mission and physical size requirements were relaxed. The result of this second effort is denoted as configuration L005 in this report; details for both designs follow.

17 A. Configuration A008 Configuration A008 was an effort to develop an aerial target of a size similar to that of the F-16 but with low observable (LO) radio frequency (RF) characteristics. This configuration was powered by two Williams FJ44-3A turbofans with a bypass ratio of 1.9 and rated at a thrust level of 3,055 lbs (each) at sea level static. The STS A008 configuration, depicted in Figure 21, was a twin engine aircraft with diamond planforms for its lifting surfaces. Leading edges are swept back at 30 degrees, while trailing edges are swept forward at 30 degrees. Figure 21. STS A008 External Geometry Basic aircraft parameters were: Wing Reference Area, Sref = sq. ft. Wing Span, b = 30.0 ft. Aspect Ratio, AR = Mean Aerodynamic Chord, MAC = ft. Nominal Take-Off Gross Weight, TOGW = 10,729 lbs. Thrust-to-Weight Ratio, T/W = 0.57

18 Figures 22 and 23 compare front and planform views of the STS A008 configuration to the F-16. Figure 22. Front View Comparison between STS A008 and the F-16 Figure 23. Plan View Comparison between STS A008 and the F-16 Provisions were made to carry 2,000 lbs of fuel in the wings and 400 lbs of fuel in the fuselage. If additional fuel were needed, it would be carried in the fuselage, where the required volume could be accommodated quite easily. There was sufficient volume in the fuselage to carry the required payload and house the auxiliary power generators and other subsystems. The landing gear would be stored in the fuselage where allowance was made for its storage. This configuration is powered by two Williams International FJ44-3A turbofans. The FJ44-3A has the following characteristics: Thrust Rating (Sea Level Static, lbf) 3,055 Flat Rated Temp (º F) 72 Specific Fuel Consumption, SFC (1/hr) Overall Pressure Ratio, OPR 21.7 Air Flow (lbm/sec) 80.2 Over All Length (in.) 62.4 Over All Height (in.) 33.4 Weight (lb) 510 Max Turbine Inlet Temperature (TIT, º F) 1,610 The propulsion data was generated using an engine deck provided by Williams International and corrected for installation losses, including spillage drag. The inlet was designed to have line-of-site blockage for minimum LO characteristics. Likewise, the exhaust nozzle geometry is configured to minimize IR exposure.

19 Performance with two FJ44s was adequate for the 5GTS mission. Table 5 lists airfield performance for the A008 while Figure 24 shows the variation of maximum specific excess power (Ps) with altitude for a nominal weight of 10,729 lbs. For comparison, F-16 takeoff and landing distances are comparable but maximum Ps is almost 600% greater at sea level. Table 5. A008 Airfield Performance Field Performance Take-Off Gross Weight Take-Off Ground Distance Air Distance to 50 Total Take-Off Distance Over 50 Obstacle Landing Weight Landing Air Distance Landing Ground Distance Total Landing Distance Over 50 Obstacle 10,729 lbs 1,831 ft 573 ft 2,404 ft 10,000 lbs 1,901 ft 2,107 ft 4,008 ft A008 Max Ps 10,729 # Ps (fps) ,000 20,000 30,000 40,000 50,000 60, Z (ft) Figure 24. Specific Excess Power vs Altitude The Specific Excess Power (Ps, in ft/sec) is the ability to accelerate or climb, or a combination of both. Therefore, the Ps values shown in Figure 24 represent the maximum rate of climb that could be achieved if all the energy available is used for climbing. Based on the values shown in the chart, the following ceiling parameters are obtained:

20 Absolute Ceiling (Zero Rate of Climb) = 49,872 ft MSL Operational Ceiling (500 ft/minute, 8.33 fps) = 48,162 ft MSL The maximum Mach number at a given altitude is determined by the thrust available and drag. The minimum Mach number is determined by the C Lmax (maximum lift coefficient) at the corresponding Mach number and altitude. In this analysis, a C Lmax value of 0.8 was used for all altitudes and Mach numbers. Maximum continuous thrust was assumed. Figure 25 shows the Mach number flight envelope for a gross weight of 10,000 lbs. The maximum Mach number value of 0.96 is achieved at altitudes from 30,000 to 35,000 ft. for maximum continuous thrust FSAT STS A008 Mach Envelope W = 10,000 lbs 50,000 45,000 40,000 35,000 Z_ft 30,000 25,000 20,000 Mmin Mmax 15,000 10,000 5, Mach No. Figure 25. Flight Envelope.

21 For cruise performance, the parameters that are significant are L/D (Lift to Drag Ratio), M*L/D (Mach times L/D), and the bottom line, SAR (Specific Air Range, miles flown per pounds of fuel burned). These parameters are shown in Figures 26 through 28. FSAT STS 30,000 ft W = 10,729 lbs L/D M*L/D Mach Figure 26. Aerodynamic Cruise Efficiency Parameters at 30,000 ft FSAT STS 40,000 ft W = 10,729 lbs L/D M*L/D Mach Figure 27. Aerodynamic Cruise Efficiency Parameters at 40,000 ft

22 Specific Air Range (nm/#) FSAT STS A008 W = 10,729 lbs SAR (nm/# ,000 40, Mach Figure 28. Specific Air Range (nautical miles flown per pound of fuel) Figure 28 indicates that configuration A008 had the best cruise performance at an altitude of 40,000 ft at Mach numbers between 0.80 and But in the mission analysis, a cruise altitude of 30,000 ft was used as specified in the FSAT design mission. Figure 29 presents the A008 turn performance at an altitude of 20,000 ft and Mach = 0.85 for maximum continuous thrust and 10,000 lbs gross weight. A maximum value of about 3.4 g s can be sustained in level flight at this condition. A value of 5 sustained g s was specified in the FSAT design mission. This performance deficiency could be significantly reduced by increasing engine thrust at the expense of engine life. This was considered a very valid trade-off for aerial target application. Engine life is usually measured in thousands of hours. Aerial targets require less than one hundred hours of service life. The other turn performance requirement, sustained 3 g s at 20,000 ft and Mach = 0.8, was met with a comfortable margin. To support the cost trade study, a weight sensitivity study was first conducted based on the aerodynamic and performance data developed for the 5 th Gen FSAT STS A008. These data were considered representative of the state-of-the-art for this type of full scale aerial target with low observable RF characteristics. The study determined how the range ratio R/R0 (where R is the range of the variable configuration, and R0 is the range of the baseline) varies with take-off gross weight for a given payload (800 lbs), and how it varies with payload for a given take-off gross weight (12,000 lbs), for the specified design mission and baseline configuration (STS A008 with 2 Williams FJ44-3 engines). The payload of 800 lbs included the actual payload (600 lbs) plus the corresponding attachment and system support (200 lbs, i.e., 33% overhead ). Therefore, all payload values should be considered carrying some weight overhead, though most likely not as high as 33% for the larger payloads.

23 STS FSAT A008 Turn Performance Williams FJ44-3 Turbofan (2) Bank Angle (degrees) W = 10,000 lbs Turn Rate and Bank Angle Z = 20,000 ft ; Mach = 0.85 Ps (ft/sec) Ps (ft/sec) Rate of Turn (degs/sec) Load Factor (g's) Figure 29. Turn Performance at 20,000 ft and Mach = 0.85 STS conducted a weight sensitivity study to quantify the effect of inaccurate estimates of airframe and payload weights. When varying take-off gross weight at constant payload, the engines were rubberized. In other words engines were scaled according to the thrust requirements so they were not real engines, but represented the current technology level. Range values show how far the aircraft would fly in optimum cruise mode burning the amount of fuel it normally would when performing the design mission. As such, they are meaningful figures of merit. Figure 30 shows the variation of the range ratio (R/R0) with take-off gross weight (GW) in pounds with the aircraft carrying a constant payload of 800 pounds. R is the range for the variable configuration while R0 is the reference range for the baseline. Figure 31 shows how the range ratio changes with payload weight at a constant take-off gross weight (12,000 lbs in this case). This relationship is linear because fuel weight is traded directly for payload. Although GW (Gross Weight) is one of the prime parameters in conceptual design, Empty Weight (WE) is a key parameter in determining aircraft cost. Therefore, the Figure 32 shows the R/R0 ratio versus WE. B. Configuration L005 The Weight Sensitivity Study presented above assumed rubberizing the engines (i.e., scaling the engines to optimum thrust requirements). However, no such engine currently exists, and one of the premises of the 5 th Gen Target Study was that new engine development was not an option. Therefore, to obtain a more realistic estimate of aircraft performance for a reduced weight configuration, the weight study was repeated with real engines. In this case, two Williams FJ44-3 engines were retained since they were still the best available selection for the reduced mission. The resulting aircraft configuration was dubbed L005 for purposes of this report.

24 R/R0 vs GW w/pl = 800 lbs R/R ,000 7,000 8,000 9,000 10,000 11,000 12,000 GW (lbs) Figure 30. Range Ratio vs Take-off Gross Weight at Constant Payload R/R0 vs GW = 12,000 lbs Range Ratio Payload (lbs) Figure 31. Range Ratio vs Payload at Constant Take-off Gross Weight

25 R/R0 vs WE (PL = 800 lbs) R/R ,000 4,000 5,000 6,000 7,000 8,000 9,000 WE (Empty Weight) lbs Figure 32. Range Ratio vs Empty Weight at Constant Payload Work on this configuration was based on the following three premises: 1. L005 aerodynamic and weight characteristics were based on those developed for STS configuration A008. Consequently, the geometry of L005 remains similar to A008 (diamond wing planform with AR = and leading and trailing edge sweeps of +-30 degrees). Baseline A008 TOGW is 10,729 lbs with a wing area of S = 320 sq. ft. yielding a wing loading of 33.5 psf. A nominal wing loading of 40.0 psf, however, was used to size L005 since, when the baseline configuration is sized to meet the original design mission, its TOGW goes up from 10,729 lbs to 12,000 lbs. The corresponding wing loading is therefore 37.5 psf. 2. As mentioned, actual engines (2 x Williams FJ44-3) were selected vice theoretical, rubber engines. Data was based on performance predicted by the engine deck provided by Williams International. 3. The original design mission profile was reduced to save weight/cost. This was possible since no official mission had been defined; therefore it was not a hard requirement. The original mission was simply a best guess developed by consensus among contractors and the USAF (AFRL and USAFA) for the 5 th Gen Target Study. The new mission was selected to achieve significant weight reduction while still maintaining operational effectiveness. For convenience, this mission was called the 80% Mission since when the amount of fuel burned for the original mission is converted into an equivalent Breguet range, the amount of fuel burned for the new mission yields 80% of the Breguet range of the original. This, of course, doesn t mean that all mission tasks were performed at 80% of the original mission. Some tasks were performed at less than 80%, others at more than 80%. Major differences in the mission profiles were: 1. Most loiter times are reduced from 15 minutes to 10 minutes 2. The high g turns are reduced to 2 turns instead of M = 0.8 and M = The Mach number for all test runs are kept at 0.90 (in the original profile they were 0.90, 0.95, and 0.95, respectively) 4. The distance flown for each test run is reduced from 50 nmi to 40 nmi (not shown in the above table) 5. The loiter after the last test run is eliminated

26 Aircraft parameters for the original mission are those of configuration A008: Wing Area, Sw = 320 sqft AR = Span = 30.0 ft TOGW = 10,729 lbs Payload WP = 800 lbs (includes 200 lbs of support/aperture structure) Fuel Weight, WF = 2,400 lbs When Configuration A008 was equipped for the original mission, i.e. actual fuel weight required to complete the mission was added in, the following results were obtained: TOGW = 12,000 lbs Fuel Weight, WF = 3,108 lbs Flight Time = hours Distance Flown = nautical miles Geometry being kept constant, the wing loading goes up from 33.5 to 37.5 psf In addition to using the 80% mission to downsize the aircraft, the payload weight was reduced to 700 lbs. The original payload was 800 lbs: 600 lbs of real payload and 200 lbs of support/aperture weight. This was probably too conservative, so 600 lbs of real payload and 100 lbs of support/aperture weight were used in revision. Requirements to complete the 80% mission and resized aircraft weight then became: TOGW = 8,800 lbs Empty Weight, WE = 6,100 lbs Fuel Weight = 2,000 lbs Fuel Burned = 1,885 lbs Flight Time = hours Distance Flown = 421 nm The geometry of the resized aircraft (L005) became: 1. Aspect Ratio, AR = Wing Loading, ω = 40.0 psf 3. Wing Area, Sw = 220 sqf 4. Span, b = 25.0 ft All other surfaces (horizontals and verticals) were reduced in area by the ratio of 220/320 = An estimated empty weight savings of nearly 1,000 lbs was achieved during this exercise, clearly showing potential for cost savings if size and mission capability were tradeable. This exercise also showed the importance of carefully defining mission requirements; an exercise that is especially difficult for targets since they have no typical mission profile. VI. Summary The configurations described above were presented to OSD management for a down-select to viable aircraft to carry forward to the Phase II Design Study. This study is discussed in the following paper. Figure 33 shows all the configurations in a common scale. The L005 and the AFRL Single-engine were both deemed too small to be considered as a Full-Scale Target and were thereby dropped from further study. The ASC design was handicapped by the required engine selection criteria. This drove the design to be too large, and therefore too expensive to justify further refinement. However, it would have been interesting to complete another design cycle using an afterburning F404, but since those engines were deemed unavailable, it would purely be an intellectual exercise. Likewise, the F125-powered delta wing design, though it promised performance sufficient to replace the QF-16, was deemed too expensive. As a result of these and previous studies, further consideration of creating a stealthy drone to completely replace the QF-16 was dropped. This left the USAFA D300 with FJ44s, the STS A008, and the AFRL Twin-Engine configurations as the only designs to be carried forward for additional analysis and testing.

27 Figure 33. Phase I Design Concepts Comparison VII. Acknowledgements The authors wish to express their appreciation to the specialists in ASC/XR who set aside pressing work to assist in this time critical study.. VIII. References Burris, P.F., Carter, D.L., White, T.L., Hayes, R.S., Brandt, S.A., and Ladd, A.F., Fifth Generation Target Study, Phase I,. AFRL-RB-WP-TR , Feb 2011.

28 Carter, D., Burris, P., Brandt, S., Fifth-Generation Target Drone Project Initial Development, Aviation Technology, Integration and Operations Conference, American Institute of Aeronautics and Astronautics, Reston, VA (submitted for publication) Brandt, S., Burris, P., Carter, D., Fifth-Generation Target Drone Phase II Design, Aviation Technology, Integration and Operations Conference, American Institute of Aeronautics and Astronautics, Reston, VA (submitted for publication) Burris, P., Brandt, S., Carter, D., Fifth-Generation Target Drone Phase III Design, Aviation Technology, Integration and Operations Conference, American Institute of Aeronautics and Astronautics, Reston, VA (submitted for publication)